This invention relates to a gas turbine engine. More particularly this invention is concerned with the design of aerofoils for gas turbine engines and in particular turbine blades or nozzle guide vanes.
An important consideration at the design stage of a gas turbine engine is the need to ensure that certain parts of the engine do not absorb heat to an extent that is detrimental to their safe operation. One principal area of the engine where this consideration is of particular importance is the turbine.
High thermal efficiency of a gas turbine engine is dependent on high turbine entry temperatures which are limited by the turbine blade and nozzle guide vane materials. Continuous cooling of these components allows their environmental operating temperatures to exceed the material""s melting point without affecting blade and vane integrity.
There have been numerous previous methods of turbine vane and turbine blade cooling. The use of internal cooling, external film cooling and holes or passageways providing impingement cooling are now common in the design of both turbines and combustors.
The shape of a nozzle guide vane or a turbine vane can substantially affect the efficiency of the turbine. The hot gases flowing over the surface of a turbine blade or nozzle guide vane forms a boundary layer around both the pressure side and suction side of the blade or vane. Ideally these flows should meet at the trailing edge of the vane causing pressure recovery and limiting the losses to friction ones only. In practice, however, the boundary layers lose energy and fail to efficiently rejoin at the trailing edge, separating and causing drag and trailing edge losses in addition to the friction losses. In order to limit these losses and improve the aerodynamic efficiency of the aerofoil it is desirable to manufacture the trailing edge as thin as possible.
However it is now essential to provide turbine blades and nozzle guide vanes with cooling holes or slots to provide both impingement cooling, internal cooling and film cooling of the blades or vanes. The blades and vanes are hollow and the internal cavities receive cooling air, usually from the compressor, which is exhausted through slots or holes at the trailing edge region.
It is known to provide the trailing edge portion aerofoils with xe2x80x98letterbox slotsxe2x80x99 through which cooling air is exhausted. The xe2x80x98letterbox slotxe2x80x99 is formed by extending the suction side of the aerofoil beyond the pressure side so as to form an overhang portion. This allows the extremity of the trailing edge portion to be thinner, hence improving aerodynamic efficiency. However there are problem with overheating and cracking of the xe2x80x98overhangxe2x80x99 portion of the trailing edge due to poor cooling thereof.
Although it is desirable to have as thin a trailing edge as possible without the need for a xe2x80x98letterbox slotxe2x80x99 arrangement, it is difficult to manufacture holes in a very thin trailing edge. There is a high scrap rate in the manufacture of such trailing edges due to the difficulty of forming holes therein. It is an aim of this invention to alleviate the difficulties associated with manufacturing trailing edges formed with cooling holes without compromising the aerodynamic efficiency of the turbine aerofoils.
According to the present invention there is provided an aerofoil member comprising a pressure surface, a suction surface, and a trailing edge portion, said aerofoil member further comprising at least one internal cavity for receiving cooling air and at least one aperture formed in its trailing edge region for exhausting cooling air from said at least one internal cavity, wherein said pressure surface is tapered toward said suction surface at the trailing edge and adjacent said aperture so as to reduce the thickness of the aerofoil member in that region.
Preferably the tapered region of said pressure surface comprises a curved portion.
Preferably the aerofoil comprises a plurality of apertures are provided in the trailing edge.